The present invention relates to gas turbine engines and, more particularly, to methods and apparatus for cooling endwalls of turbine airfoils.
In a gas turbine engine, hot gas exits a combustor and is utilized by a turbine for conversion to mechanical energy. This mechanical energy drives an upstream high pressure compressor. The turbine comprises a plurality of rows of blades which are carried by a turbine rotor, alternating with rows of stationary nozzles. The turbine blades and nozzles are subjected to a flow of the corrosive, high-temperature combustion gases. These “hot section” components are typically cooled by a flow of relatively low-temperature coolant, such as air extracted (bled) from the compressor.
As turbine inlet temperatures in modern gas turbine engines continue to rise, the endwalls of the hot section components (i.e. turbine blade platforms and nozzle bands) become more difficult to cool with traditional techniques. In addition, advanced aerodynamic features such as endwall contouring put extra pressure on maintaining acceptable material temperatures.
The current state of the art is to drill film holes through the endwalls, to be fed by cooling air beneath the component. As a result, holes can only be placed in certain regions where they can be completely drilled to the other side or where the gas path pressure is low enough since the cooling air pressure feeding these holes is much lower than the airfoil cooling air.
Some designs use hollow platforms that feed compressor bleed air to film cooling holes, but these designs are generally not adaptable to providing different cooling hole patterns based on varying operating conditions.
Accordingly, there is a need for a turbine airfoil platform with improved cooling.